Hybrid rocket engine and method of propelling a rocket

ABSTRACT

A hybrid rocket engine and a method for propelling a rocket utilizing a vortex flow field. The flow field includes an outer fluid vortex spiraling toward a closed end of the flow field generating apparatus and an inner fluid vortex substantially concentric with the outer vortex spiraling away from the closed end and toward an outlet opening in which the inner vortex spirals in the same direction as the outer vortex, but in the opposite axial direction. The invention also relates to a rocket propulsion system utilizing the flow field in which the propulsion system includes a combustion chamber with a fuel source and an oxidizer source that deliver the said fuel and said oxidizer to the said outer and inner vortexes in a manner to support a combustion process while flowing along the flow field.

This application is a continuation-in-part of U.S. application Ser. No.09/337,457 filed Jun. 21, 1999, (now U.S. Pat. No. 6,298,659) whichclaims the benefit of Provisional Application Ser. No. 60/125,910 filedMar. 24, 1999.

BACKGROUND OF THE INVENTION

1. Field of the Art

The present invention relates generally to a vortex flow field and theapparatus and method to produce and sustain it and more particularly toa hybrid rocket engine and a method of propelling a rocket utilizingsuch vortex flow field. The flow field in accordance with the presentinvention is capable of providing separate regions or zones within andamong one or more flowing fluids contained within a common chamber,without the need for diaphragms or other physical separators orbarriers. It is evident and believed that the flow field of the presentinvention has utility to a wide range of applications. One general fieldof application is that of combustion chambers, and more particularly,that of combustion chambers and methods for rocket engines or the likeand hybrid rocket propulsion systems. A combustion chamber and method inaccordance with one embodiment of the present invention utilizes theunique flow field of the present invention to improve hybrid rocket fuelregression and increase mixing length in a rocket or other engine.Another embodiment is in the form of liquid rocket engine to prevent hotcombustion products from contacting the chamber wall.

2. Description of the Prior Art

Virtually countless applications exist for a flow field which is compactand is capable of providing one or more separate regions or zones offlowing fluids within a container, without substantial mixing andwithout the need for any physical barrier or other separators betweensuch regions or zones. With such a flow field, a chemical reaction, suchas combustion, can be induced to incur in one region or zone while aseparate fluid or process occupies another region or zone.

Many devices depend upon vortex flows for their successful operation,such as combustion chambers, cyclone separators, classifiers and thelike that are in common use. All of these devices introduce swirlingflow at one end of a passageway in which the flow follows a generallyhelical path to exit at the opposite end. Such conventional vortex flowsdo not achieve zonal separation as does the unique flow field that isthe subject of the present invention.

Although the flow field in accordance with the present invention hassignificant applications in a variety of fields, it has particularapplication to the field of rocket engines and in one embodiment,specifically to hybrid rocket engines. Hybrid rocket engines denote aclass of rocket propulsion systems in which one propellant is in fluidform and the other propellant is in the form of a solid grain.Typically, the fluid propellant is the oxidizer and the solid grain isthe fuel. The oxidizer such as liquid oxygen is sprayed into thecombustion ports in the solid fuel grain and caused to ignite. The hotcombustion products sustain the combustion process until either theoxidizer flow is shut off or the fuel grain is depleted. In virtuallyall contemporary hybrids of today the limiting design factor is the rateat which the solid fuel grain can be caused to burn. The burn rate,often expressed as regression rate, is the rate at which fuel can beinduced to vaporize or ablate off the grain surface so it canparticipate in the combustion process and contribute to rocket thrust.Because the rate is typically slow, conventional hybrid fuel grains mustbe made with large exposed surface areas. This is accomplished bycasting large open combustion ports in the grain. The large ports wastevolume in the high pressure casing, so that a larger, heavier, and moreexpensive case is needed than would be required if the fuel graincombustion ports could be much smaller by means of a flow field whichimproves the regression rate.

In recent years, hybrid rockets have received increasing attention fromthe National Aeronautics and Space Administration (NASA) sectors,Department of Defense, industrial aerospace participants and researchinstitutions because their unique operational characteristics arecapable of providing safer, lower-cost avenues to space thanconventional solid propellant and liquid bi-propellant rocket propulsionsystems. For example, hybrid rocket engines can be easily throttled forthrust tailoring, to perform in-flight motor shutdown and restart and toincorporate non-destructive mission abort modes. Also, since fuel in ahybrid rocket engine is stored in the form of a solid grain, suchengines require only half the feed system hardware of liquidbi-propellant engines. Still further, the commonly used butadiene-basedsolid grain fuels are benign and neither toxic nor hazardous for storageand transportation. The hybrid solid fuel grain is also not susceptibleto cracks and imperfections that can lead to catastrophic failure insolid rocket motor propellant grains.

However, despite these benefits, classical hybrid rocket engines, inwhich the oxidizer gas is injected into the combustion chamber at theend opposite the exit nozzle and in a direction parallel to the solidfuel grain, have not yet found widespread use for either commercial ormilitary applications. Reasons for this include the fact that theysuffer from relatively slow solid fuel regression rates, low volumetricloading and relatively poor combustion efficiency. For example,polymeric hybrid fuels such as hydroxyl-terminated polybutadiene (HTPD)regress generally about an order of magnitude slower than solid rocketmotor propellants. In an effort to overcome these lower regressionrates, complex cross-sectional geometries of the hybrid solid grain fuelwith large wetted surface areas are often employed to achieve a largemass of flow rate of pyrolyzed vapor from the fuel grain. It has beenshown that a three to fourfold increase in fuel regression rate canresult in significant cost reductions, simplified grain manufacturingand large reductions in rocket inert weight.

In addition to problems associated with the low regression (fuelburning) rates of hybrid engines, the short straight line travel of thepyrolyzed fuel grain vapor and oxidizer as they traverse the combustionregion results in incomplete mixing. This often necessitates the use ofsecondary combustion chambers at the end of the fuel grain to completethe combustion process. These secondary chambers add length and weightto the overall design and have the additional disadvantage of serving asa potential source and location of combustion instability.

Furthermore, both conventional hybrids and solid rocket motors mustprovide insulation layers between the solid propellant grain and thehigh pressure casing wall. This is necessary to prevent the exposure ofthe casing to the high temperature combustion gases when the grainmaterial has been burned away out to the casing and no longer providesprotection. The insulation adds weight and cost to the motor.

Accordingly, there is a need in the art for a flow field, and astructure and method for producing and sustaining it, which providesseparate regions or zones of flowing fluids within a chamber. There isalso a need in the art for a combustion chamber and method utilizingsuch a flow field, and particularly a combustion chamber and method fora hybrid rocket engine, which significantly increases the regressionrate of the solid fuel grain and the effective chamber length and mixingwithin the combustion chamber. There is also a need for a combustionchamber and method utilizing such a flow field that prevents the hotcombustion products from reaching the chamber wall.

SUMMARY OF THE INVENTION

In accordance with the present invention, a fluid flow field, and astructure and method for producing and sustaining the field, has beendesigned. This flow field provides for separate regions or zones offlowing fluids within a chamber without the need for physical barriersor other separators and without substantial mixing between the regionsor zones.

In a revolutionary departure from prior art the present inventionintroduces the incoming swirling flow concentric to the outlet passageand by this means establishes a new and unique flow field nothere-to-fore known or described in literature, nor does it have anypreviously known existing physical embodiments beyond those defined anddescribed herein. The flow field inherently divides into an outerupwardly flowing vortical helix along a chamber wall, an inner downwardflowing vortical helix along the center region of the chamber, aconverging flow field at the head end where the outer vortex transformsinto the inner vortex, a converging flow field as the flow approachesthe exit nozzle, and less well defined regions of velocities andpressure gradients elsewhere throughout the chamber.

The distinct regions can be controlled by chamber geometry, fluidinjection parameters, external heat addition, and combustion or otherchemical reactions to produce certain desired and specific results.These reactions include, but are not limited to, enhanced combustion ofmaterials forming the chamber walls, limitation of combustion to occurin the center vortex only, combination of reactions at the wall, withsubsequent separate and different reactions in the central vortex, andfluid distillation and liquid-vapor separation.

The flow field is produced by injecting flow tangentially into acylindrical chamber which is substantially closed at one end and whichhas a converging outlet at the other end. In the preferred embodiment,the flow is introduced into the interior of the chamber near the outletend of the chamber and in a direction which is substantially tangent to,or which creates a flow which is substantially tangent to, the innerwall of the chamber. This tangential injection causes the flow in thechamber to swirl and follow a spiral path up the inner wall of thechamber thereby establishing an annular vortex flow bounded by the innerwall of the chamber. When the spiral flow reaches the closed end of thechamber, the flow inherently translates inwardly to the center of thechamber to form the second vortex where the flow moves spirally awayfrom the closed end, down the core of the chamber and out the chambernozzle. This inner vortex or spiral flow through the center of thechamber rotates in the same direction as the outer vortex, but at asmaller radius and thus a greater angular velocity in accordance withthe principle of angular momentum conservation. The result of the aboveis the establishment of a radial pressure gradient field throughout thechamber. One key feature of this pressure gradient field occurs at theexit nozzle. Specifically, pressure at the nozzle converging wallincreases and pressure at the swirl axis decreases. Accordingly,injection flow at the periphery of the vortex near the outlet end andtangential to the outer vortex streamline cannot penetrate the pressuregradient that has formed by the inner vortex at the nozzle convergingregion. Thus, this incoming flow cannot flow toward the exit. Instead,it must take an alternate flow path to enter the lower pressure regionin the center of the vortex flow approaching the converging nozzlesection. This alternate path is up along the wall and then inward to thecenter vortex before flowing down and out the nozzle. Accordingly, asthe inner vortex flow approaches the nozzle, it enters the convergingsection of the nozzle, thereby increasing the swirl or angular velocityand thus producing an enhanced radial pressure gradient that blocks theoutflow of the fresh incoming stream.

The above-described flow field has several unique characteristics.First, the flow path of the injected fluid before reaching the outlet isquite long and highly convoluted. Thus, it provides an opportunity forintense and extensive mixing along the flow path, particularly in thecore or inner vortex where the angular velocity of the swirl is greater.Secondly, the outer and inner vortexes are individually discrete. Thus,the fluid flow in the inner vortex does not mix significantly with thefluid flow in the outer vortex. This enables the inner vortex to supportburning or other chemical reactions to some significant degreeindependent of the outer vortex. Because of this, materials such aspropellant or other chemicals, can be added to the inner vortex byinjection at the conjunction of the two vortices at the closed end ofthe chamber and cause combustion or other chemical reaction to occur andbe sustained wholly in the inner vortex if so desired.

The ability to produce and sustain the above-described double vortexfield flow has countless potential applications and several immediatepractical applications. By way of example only, one immediate practicalapplication of the flow field of the present invention is in the fieldof rocket propulsion.

In such an application, utilization of the flow field of the presentinvention facilitates a combustion chamber and method which providesdramatically increased regression rates of the fuel grain and increasedmixing length and improved mixing within the combustion chamber. In apreferred embodiment and application, the present invention provides fora combustion chamber and method for use in a hybrid rocket engine.

In applying the double vortex flow field of the present invention to thepreferred embodiment of a hybrid rocket engine, the flow is created byinjecting one component of the combustion mixture (such as the oxidizer)into a generally cylindrical combustion chamber which is closed at oneend and is provided with a converging outlet nozzle at the other end. Byinjecting the flow of oxidizer fluid in a direction circumferentiallytangent to the inner wall, the fluid is caused to swirl and advance upthe cylinder wall in a vortex pattern toward the closed end. When thisouter vortex flow reaches the closed end, it moves radially toward thecenter of the chamber and continues to move in a swirling vortex alongthe middle or core of the chamber and out through the exit nozzle. Ifthe inner walls of the chamber are hybrid fuel grain and the fuelgrain/oxidizer combination is ignited, several unique and advantageouscharacteristics result. First, the high velocity outer vortex scrubs theburning fuel grain surface, causing enhanced heat transfer to thesurface. Combustion near and on the surface is also able proceed becausefresh oxidizer is carried directly to the surface by turbulent transportmechanisms in addition to the usual molecular diffusion process. Second,the vortex also sustains radial pressure and density gradients thatcause hot, low density combustion products to be buoyed out of thecombustion zone so their presence does not hinder the combustionprocess.

Third, because the flow path of the injected fluid (the oxidizer) toreach the outlet is very long and highly convoluted, it provides anopportunity for intense and extensive mixing and combustion with thefuel grain vapor, particularly in the core or inner vortex. Accordingly,in the above application, the outer vortex flow causes rapid burning ofthe fuel grain on the wall of the cylinder, and the inner vortex causescombustion to proceed rapidly, by providing intense mixing andcombustion travel distance to allow combustion to reach completion,thereby achieving high combustion efficiency.

In general, the structure to produce the flow field of the presentinvention as well as the structure of the combustion chamber inaccordance with the present invention includes a containment chamberwith first and second ends which are sometimes referred to as head andtail ends or closed and nozzle ends. In the preferred embodiment, thecontainer inner wall is covered with a solid fuel grain or a fuelsource. The chamber is closed at one end and provided with an exitnozzle at its opposite end. One or more fluid (oxidizer) delivery portsare provided near the end of the container adjacent to the nozzle forthe purpose of delivering an oxidizer (or other fluid) into the chambertangentially to the inner surface of the fuel grain coating the innerwall of the chamber. After injection, the oxidizer swirls along thesurface of the fuel grain toward the closed end, at which location itmoves radially toward the center and then swirls in the form of theinner vortex toward the nozzle end of the cylinder.

The method aspect of producing the flow field of the present inventionincludes providing a cylindrical chamber with a closed head end and anopposing nozzle end and injecting a fluid tangentially to the innerwall. In the preferred application of the present invention thecylindrical chamber is a combustion cylinder with a closed end and anozzle end and the inner surface of the chamber is provided with a fuelsource. The injected fluid is an oxidizer. Upon injection of theoxidizer, the oxidizer/fuel mixture is ignited.

Accordingly, it is an object of the present invention to provide a flowfield, and a structure and method for producing and sustaining such flowfield which provides distinct and separate regions of flowing fluidwithin a chamber, without the use of physical barriers.

Another object of the present invention is to provide an improvedcombustion chamber and method utilizing the above-described doublevortex flow field.

Another object of the present invention is to provide an improvedcombustion chamber and method utilizing the above flow field and toprovide for increased fuel regression rates and increased traveldistance and mixing to achieve complete combustion.

A further object of the present invention is to provide a hybrid rocketengine utilizing the above-described double vortex flow field.

A further object of the present invention is to provide a liquid rocketengine utilizing the above-described vortex flow field.

A still further object of the present invention is to provide animproved hybrid rocket propulsion system that facilitates and promoteshigh and uniform fuel grain regression rates so that small combustionports can be used in the propellant solid grain.

Another object of the present invention is to provide a hybridpropulsion system that inherently cools the case walls whenever fuel isnot present to insulate the wall from hot combustion products.

A more specific object of the present invention is to provide a hybridrocket propulsion system that creates and uses a unique internalcombustion vortex flow field to enhance grain regression rate and toincrease the efficiency of the combustion process.

Another object of the present invention is to provide a combusting flowfield that allows the use of a single grain port for the combustionprocess.

A further object of the present invention is to provide an injectionmeans for the fluid propellant that induces the double vortex flow fieldin the grain combustion port.

A further object of the present invention is to provide a combustionprocess that inhibits combustion instability.

Another object of the present invention is to provide a double helixflow field in which an outer helix flows upwards along the grain surfaceinducing combustion, and an inner combustion helix flows down the portcenterline and out the nozzle to produce thrust.

To achieve the foregoing and other objects and in accordance with thepurpose of the present invention, a self-contained propulsion system isprovided with a motor casing that houses a solid propellant grain. Afirst fluid propellant that will combust when in the presence of thesolid propellant in the presence of an ignition source, is storedseparately from the solid propellant in a fluid tank. A delivery meanssupplies at least a portion of the said fluid propellant in eitherliquid or gaseous state to the combustion port of the solid grain. Anignition means initiates combustion with the combustion port of thesolid propellant grain. A fluid injection means that will cause thefluid propellant to enter the solid propellant grain case in such amanner as to form an upflowing helix along the surface of the combustionport in the solid propellant grain and then a downflowing helix alongthe centerline of the combustion port, said downflowing helix toeventually exit the chamber via the discharge nozzle.

The fluid propellant can be provided to the entrance to the fuel graincase by any of various common means, including delivery from pressurizedtanks, or by pumps of suitable designs. The fluid can be either theliquid or gaseous state. Commonly the fluid propellant is the oxidant.In one embodiment, the oxidant is burned in a highly oxidizer-richcombustor (termed a “preburner”) and the resulting oxidizer-richcombustion products are used to drive a turbopump that pressurizes theliquid oxidizer for delivery to the preburner. After driving theturbine, the oxidizer-rich combustion products leave the turbine andflow to the injection ports of the fuel grain high pressure casing. Theoxidant enters the fuel grain ports in a fluid phase that may be at highenough pressure to be supercritical. The injector elements arepositioned and designed such that the injected flow develops theco-axial vortex flow field within the chamber in the manner that is thesubject of this invention.

In another embodiment, the oxidizer in liquid state is carried in a highpressure tank. Pressurant is supplied by a conventional tankpressurization system well known to those acquainted with theprofession. The liquid oxidant is expelled from the tank and deliveredat high pressure to the injection ports of the fuel grain high pressurecasing. The oxidant enters the casing in the liquid state and is quicklyheated and vaporized as it enters the combustion port of the fuel grain.The injector elements, in concert with the cylindrical casing andcylindrical combustion port in the fuel grain, are designed to impart astrong swirl component to the injected oxidant flow. The swirl acts inthe chamber to develop the co-axial vortex flow field that is a keyaspect of this invention.

Typical oxidants are oxygen in liquid or gaseous form, inhibited redfuming nitric acid (IRFNA), hydrogen peroxide, nitrogen tetroxide andnitrous oxide.

The solid fuel grain may be of any suitable material. A preferredmaterial used for hybrid fuel is hydroxyl-terminated polybutadine, acomplex, rubber-like hydrocarbon formulation that is readily cast andcured at modest cost. Other fuels include paraffins and PMMA. Fillerssuch as aluminum powder and boron may be added to customize performance.The propellants chosen are not critical to the technology of thisinvention.

The above and other objects of the present invention will becomeapparent with reference to the drawings, the description of thepreferred embodiment and the appended claims.

DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic representation of the flow field of the presentinvention and the apparatus for producing and sustaining the same.

FIG. 2 is a cross sectional view of the apparatus of FIG. 1 showing thefluid inlet passages for creating the flow field of the presentinvention.

FIG. 3 is one embodiment of a test device for a propulsion system for aliquid rocket engine utilizing the flow field of the present invention.

FIG. 4 is a further embodiment of a test device for a propulsion systemfor a liquid rocket engine utilizing the flow field of the presentinvention.

FIG. 5 is a further embodiment of a propulsion system for a hybridliquid rocket engine utilizing the flow field of the present invention.

FIG. 6 is a further embodiment of a propulsion system for a hybridliquid rocket engine utilizing the flow field of the present invention.

FIG. 7 is a schematic illustration of a two stage rocket includingpropulsion systems utilizing the flow field of the present invention.

FIG. 8 is a further embodiment of a propulsion system for a rocketengine.

FIG. 9 is a further embodiment of a propulsion system for a rocketengine.

FIG. 10 is a view, partially in section, as viewed along the sectionline 10—10 of FIG. 9.

DESCRIPTION OF THE PREFERRED EMBODIMENT

The flow field of the present invention is a double helical vortex flowfield comprised of an outer vortex and a substantially concentric innervortex which rotates in the same direction, but at a higher angularvelocity. The inner and outer vortices are substantially discreet andadvance in opposite axial directions. As described in greater detailbelow, the flow field of the present invention comprises a flow fieldaxis having a first end and an opposite constricted second or exit endand three separate and discrete flow regions. These include a first flowregion comprising a first or outer, substantially vortex fluid flowaround the flow field axis and toward the first end, a second flowregion comprising a substantially radial fluid flow from the firstvortex fluid flow at the first end toward the flow field axis and athird flow region comprising a second or inner substantially vortexfluid flow around the flow field axis from the first end and toward thesecond end, with the second vortex flow field being radially inwardly ofthe first vortex fluid flow. The flow field of the present invention hasa wide variety of potential applications. One application is as acombustion chamber for a propulsion system and more specifically as acombustion chamber for a hybrid rocket propulsion system. A preferredembodiment of the flow field of the present invention will be describedwith reference to the schematic representation of the flow field and theapparatus for producing and sustaining the same as shown in FIG. 1 andwith reference to various embodiments of a propulsion system for ahybrid rocket engine utilizing such a flow field.

With reference to FIG. 1, the apparatus for producing and sustaining theflow field of the present invention includes a containment cylinder 10comprised of a generally cylindrical side wall 11 with an inner wallsurface 12. The containment cylinder 10 defines a hollow interior flowfield chamber 14 having a chamber axis 15 extending generally parallelto the side wall 11. The side wall 11 further includes a first end 16and a second end 18 opposite the first end 16.

The containment cylinder 10 includes a closed end defined by the closedend wall 19 which extends from the first end 16 of the side wall 11 andis intersected by the chamber axis 15. The opposite end of thecontainment cylinder 10 is an outlet opening end defined by the wallportion 20. The wall portion 20 extends from the second end 18 of theside wall 11 and includes a restricted opening 21. In the preferredembodiment and throughout the application, the opening 21 may bereferred to as a restricted opening in the sense that its diametricaldimension is less than the diametrical dimension of the inner wallsurface 12 and thereby restricts the flow of fluid from the flow fieldchamber 14 out through the opening 21.

The side wall 11 includes one or more fluid inlets or inlet openings 22positioned about the periphery of the side wall 11 and extending intothe chamber 14. As shown best in FIG. 2, these inlet openings 22 extendthrough the side wall 11 so that they are directed substantially tangentto the inner wall surface 12. Although any number of openings 22 may besufficient to produce and maintain the flow field of the presentinvention, it is preferable for there to be at least two or more inlets22. In the preferred embodiment, as best shown in FIG. 2, all of theinlets 22 are directed in the same angular direction relative to thechamber axis.

The inlets 22 can be positioned at multiple axial and radial positionsin the side wall 11. In the preferred embodiment, however, the inlets 22are positioned at the same axial position near the restricted openingend 18 of the side wall 11. This enables the incoming fluid to beginforming the flow field near the restricted opening end of the cylinder10 to achieve maximum benefits from the flow field.

In the preferred embodiment, the containment cylinder 10 is shown as asubstantially right cylindrical member meaning that the cylinder 10 hasa circular cross-section and the radial distances from the axis 15 toall points on the inner wall surface 12 are equal. However, variousother shapes of containment cylinders or containment vessels can beutilized as well. For example, it is contemplated that the side walls 11could be converging or diverging at any point along their length as theyextend between the ends 16 and 18. An example of such a structure wouldbe a frusto-conical configuration. The side walls 11 could also becurved as they extend from one end 16 to the other 18. For example, theside wall as it extends from one end to the other could have a singleradius of curvature resulting in the containment vessel or cylinderhaving a generally spherical configuration. Or the side wall, as itextends from one end to the other, could have multiple curvaturesresulting in elliptically shaped or other more complex configurations.Preferably the containment cylinder or vessel 10 has a side wall whichis a revolute surface in which the inner wall surface 12 is defined bythe revolution of a line (straight or curved) relative to the chamberaxis 15. In other words, all points on the inner surface 12 of the sidewall 11 which intersect a plane perpendicular to the axis 15 areequidistant from such axis.

The flow field in accordance with the present invention is formed byintroducing a fluid into the chamber 14 through the inlet openings 22.Upon being introduced into the interior of the chamber 14 through theinlets 22 in a direction generally tangential to the inner wall surface12, the fluid flow under its injected momentum, is forced by the innerwall surface 12 to swirl around the surface 12 in a spiral or vortexpattern to form an outer vortex 24. As the introduction of fluidcontinues, the outer vortex 24 continues to advance up the inner wallsurface 12 toward the closed end of the chamber 14 defined by the endwall 19. During this outer vortex flow 24, a pressure gradient iscreated within the chamber 12 in which the pressure increasesexponentially proportional to the radius of the vortex flow from thechamber axis 15. Thus, the above-described outer vortex flow 24 createsa higher pressure area at the inner surface 12 where the radial distanceis greater and a pressure gradient toward the center of the chamber 14along the chamber and vortex axis 15.

When the outer vortex flow 24 reaches the closed end of the chamber 14,the flow travels radially inwardly toward the chamber and vortex axis 15and the lower pressure area and turns down and continues to swirl in aspiral or vortex pattern toward the outlet opening 21. This vortex flowforms an inner vortex 25 which is internal and co-axial of the outervortex 24, and which swirls in the same direction, but which advancesthrough the chamber 14 in the opposite axial direction. The inner vortex25 also creates a pressure gradient which increases exponentiallyproportional to the radial distance of the rotating fluid from the axis15. Further, because the vortices formed in the chamber 14 approximate a“free vortex”, angular momentum is conserved at all radii. Thus, theinner vortex 25 rotates or swirls about the axis 15 at a greater angularvelocity than the outer vortex 24. As the center vortex approaches theexit nozzle, it is forced to converge to the smaller diameter of thenozzle opening. The angular velocity of the flow at the nozzle entrancethus becomes higher and creates a high local pressure gradient at thenozzle entrance. One result of this high pressure gradient is that thelocal pressure on the nozzle entrance wall is higher than any place elsein the chamber. This local high pressure prevents significant axial flowfrom passing through this local region. Thus, new inflow injectedthrough the inlets 22 is forced by the inner vortex 25 to flow along theouter vortex flow path 24 toward the closed end wall 15 and theninwardly to the inner vortex flow path 25 toward the outlet 21. Uponreaching the outlet opening 21, the swirling fluid from the inner vortexexits the chamber 14.

This flow field has several unique characteristics. First, the flow pathof the injected fluid from the inlet 22 along the outer vortex 24, alongthe inner vortex 25 and out through the opening 21 is quite long andhighly convoluted. Thus, it provides an intense and extensive mixing ofthe injected fluid, particularly in the core or inner vortex 25 wherethe angular velocity of the swirl is greater. Secondly, the outer vortex24 and the inner vortex 25 are individually discreet. In other words,because of the various pressure and velocity gradients formed by theouter vortex 24 and inner vortex 25, the fluid flow in the inner vortex25 does not mix significantly with the fluid flow in the outer vortex.This enables the inner vortex to support burning or other chemicalreactions independent of the outer vortex 24. Because of this, reactionsinitiated in the outer vortex 24 can continue through the inner vortex25 and materials such as additional oxidizer or other chemicals can beadded to the inner vortex 25 by injection at the conjunction of the twovortices at the closed end of the chamber.

Accordingly, the apparatus for producing and sustaining the flow fieldof the present invention can be summarized as including a flow fieldchamber comprising a chamber axis extending substantially through thecenter of the chamber, a side wall with an inner surface encircling thechamber axis, a closed chamber end extending from one end of the sidewall and intersected by the chamber axis and an outlet chamber endextending from the opposite end of the side wall and intersected by thechamber axis. The outlet chamber end includes a restricted fluid outletand the chamber includes one or more inlet openings directedsubstantially tangent to the inner surface. The method of producing andsustaining the flow field of the present invention includes the step ofintroducing a fluid flow into the chamber through the inlet openings ina direction substantially tangent to the inner wall surface. It ispreferable for at least some of the inflow to be introduced through thewall inlet openings near the level of the outlet end. This produces thefield flow of the present invention which comprises an outer vortexadjacent to the inner wall and an inner vortex internal to the outervortex which rotates in the same direction as the outer vortex, butrotates at a higher angular velocity and flows axially in the oppositedirection from the outer vortex.

The above-described flow field of the present invention has countlesspotential applications where extensive fluid mixing is desired alongdiscrete flow paths or where significant surface contact between agaseous fluid and a second material is desired. One such application isfor use as a combustion chamber in the field of propulsion, and moreparticularly to a combustion chamber for use in connection with a hybridrocket propulsion system. The preferred embodiment of such anapplication is described with reference to FIGS. 3-7.

Reference is first made to FIG. 3 illustrating a propulsion system for ahybrid rocket engine. Specifically, the system of FIG. 3 includes acombustion chamber 26 utilizing the flow field of the present inventionand a solid fuel or solid fuel grain together with an oxidizer as thepropellant mixture. As shown, the combustion chamber 26 is defined by aninner cylindrical wall 28, a closed end wall 29 and a nozzle end wall30. An outer cylindrical wall 31 concentric with the inner wall 28 andspaced radially outwardly therefrom function to form a generally annularoxidizer manifold 32 between the walls 28 and 31. In the preferredembodiment, the inner wall 28 and the outer wall 31 are connected withthe closed end wall 29 and nozzle end wall 30 as shown. The nozzle endwall 30 is provided with a nozzle assembly 34 including a nozzle opening35 and a diverging nozzle outlet 36. The closed end wall 29 is providedwith one or more primary oxidizer supply ports 38, 38 from an oxidizersource for providing fluid oxidizer such as gaseous oxygen to themanifold 32. A plurality of secondary oxidizer inlets 42, 42 are alsoprovided in the wall 30 to provide additional oxidizer into thecombustion chamber 26 if desired. Such additional oxidizer may bedesired to promote and ensure full combustion of the oxidizer-fuelmixture during its passage through the flow field.

In the embodiment of FIG. 3, the inner wall 28 is constructed of a solidfuel grain such as hydroxyl-terminated polybutadiene (HTPB) providedwith a plurality of oxidizer inlets 39. Similar to the inlets 22 ofFIGS. 1 and 2, the inlets 39 of FIG. 3 are directed in a directionsubstantially tangent to the inner surface of the side wall 28. Thus,when fluid oxygen is introduced through the inlets 39, it swirls in aspiral pattern along the inner surface of the wall 28 to form an outervortex 40. This outer vortex swirls toward the closed end wall 29 andthen inwardly toward the center of the chamber 26. From here it swirlsdownwardly through the center of the chamber 26 and toward the nozzleend wall 30 in a continuing spiral pattern to form the inner vortex 41.From here, the fluid in the inner vortex flows out through the nozzleopening 35 and nozzle outlet 36.

Because the inner wall 28 is formed of a solid fuel grain, combustion orburning occurs when an oxidizer such as oxygen is injected into thechamber 26 through the inlets 39 and contacts the solid fuel grain inthe presence of ignition conditions. Specifically, as the gaseous oxygenenters the combustion chamber 26, it is forced toward the closed endwall 30 by the favorable axial pressure gradient and forced outwardly bycentrifugal acceleration resulting from the tangential direction ofinjection. The oxidizer spirals along a helical path upwardly along thefuel surface of the wall 28, mixing and burning with vaporized fuel fromthe grain surface. At the top of the chamber, this outer burning vortex40 converts into a downward spiraling inner vortex 41 that eventuallypasses through the nozzle opening 35. The combustion results inincreased temperature within the combustion chamber 26 and thusincreased angular velocity and mixing within the outer vortex 40 andinner vortex 41. As a result of the extensive mixing and highlyconvoluted flow path along the outer and inner vortices, the combustionis highly efficient. Further, because of the extensive surface contactbetween the oxidizer and the fuel grain along the outer vortex 40, thereis an increased burning or combustion rate of the solid fuel, therebysignificantly increasing the solid fuel regression rate. Tests havedemonstrated as much as an 800% or more increase in solid fuelregression rates over those of comparable classic conventional hybrids.

Although ignition can be caused to occur at any location along the innerwall 28, it preferably is caused to occur near the nozzle end of thechamber 26 and then it quickly spreads upwardly over the entire graininner surface. This results in the maximum benefit of the combustionflow field of the present invention.

The embodiment of FIG. 4 is similar to that of FIG. 3 except that theinner wall 44 is not provided with inlet openings along its length likethe inner wall 28 of FIG. 3. Instead, the wall 28 is substantially solidand cylindrical as shown. Further, the embodiment of FIG. 4 includes aplurality of oxidizer inlets near the nozzle end for introducingoxidizer into the chamber 26 and a solid fuel grain layer 46 is providedor formed on the inner surface of the wall 44 to serve as the fuel withwhich the oxidizer reacts during combustion. Otherwise the embodiment ofFIG. 4 is the same as that of FIG. 3.

During operation, oxidizer in the form of oxygen or other appropriateoxidant is introduced into the manifold 32 through the inlets 38. Theoxidizer then enters the chamber 26 through the inlets 45 and forms theflow field of the present invention with the outer vortex 40 and theinner vortex 41 as shown. During this passage along the flow field, theoxidizer contacts the fuel grain 46 and the same is ignited, preferablynear the nozzle end of the chamber. A common solid fuel 46 is HTPB,although other solid fuels of designer choice may be used as well.

FIG. 5 illustrates a further embodiment of a propulsion system for abi-propellant liquid rocket engine comprising a combustion chamberutilizing the flow field of the present invention. In FIG. 5, the systemincludes a combustion chamber 48 defined by a porous fuel injection wallcomprised of a generally cylindrical side wall portion 49 and agenerally planar end wall portion 50. In the preferred embodiment, thewall portions 49 and 50 are porous to permit passage of liquid orgaseous fuel from the fuel manifold 51 and into the chamber 48. Themanifold 51 is defined on its outer side by the chamber housing whichcomprises a cylindrical portion 52 conforming in shape and configurationto the wall portion 49 and a closed end portion 54 conformingsubstantially to and spaced outwardly from the planar wall portion 50.

The nozzle end of the propulsion system of FIG. 5 is provided with aregenerative cooling nozzle skirt 55 comprised of the skirt inner wall56, the skirt outer wall 58 and the cooling skirt chamber or passageway59 positioned between the walls 56 and 58. Preferably the skirt innerwall 56 includes a converging portion 56 a and a diverging portion 56 b.Similarly, the skirt outer wall 58 includes a converging portion 58 aand a diverging portion 58 b. The lower end of the diverging portion ofthe nozzle includes a fuel inlet 60 in communication with the coolingskirt passageway 59. As shown, the cooling skirt passageway 59 is alsoin communication with the fuel supply manifold 51. Thus, liquid orgaseous fuel provided through the fuel inlet 60 flows through thecooling skirt chamber 59 and into the fuel supply manifold 51. Fromthere it flows through the porous fuel injection wall portions 49 and 50into the combustion chamber 48.

The nozzle end of the cavity 48 is further defined by a wall portion 61positioned as an extension of the fuel supply wall 49. The wall portion61 is provided with one or more oxidizer inlets 62. Like the oxidizerinlets 22 of FIGS. 1 and 2 and the inlets 39 and 45 of FIGS. 3 and 4,the inlets 62 are disposed in a direction substantially tangent to theinner surface of the wall portions 49 and 61. A generally annularoxidizer manifold 64 defined by the manifold housing 65 extends aroundthe chamber and provides a passage for the flow of oxidizer such asoxygen through the inlet 66 into the manifold 64 and then through theinlets 62 and into the chamber 48.

Similar to the other embodiments, the introduction of the oxidizerthrough the inlets 62 creates a flow field comprising an outer vortex 68and an inner vortex 69. During the passage of oxidizer along this path,it contacts and mixes with the liquid or gaseous fuel flowing inwardlythrough the porous wall portions 49 and 50 along the paths defined bythe directional arrows 70. When ignited, the oxidizer and fuel bum tocreate the high pressure combustion products that are expelled throughthe nozzle to produce the propulsion force.

One primary feature of the embodiment of FIG. 5 is that it facilitatesthe use of liquid or gaseous fuel such as liquid Rocket Propellant 1(RP-1) and liquid or gaseous hydrogen, among other appropriate rocketpropellants which designers may choose. The embodiment of FIG. 5 alsoprovides a means for cooling the nozzle end of the system as well as theouter chamber housing walls 52 and 54 as the fuel flows through thecooling skirt 59 and the fuel supply manifold 51. This facilitatesconstruction of the rocket propulsion system of FIG. 5 from materialswhich are less heat resistant and thus less expensive than previousliquid rocket engine designs.

In the embodiment of FIG. 5, the fuel is shown as being provided throughthe porous wall portions 49 and 50, while the oxidizer is shown as beinginjected through the inlets 62. It is permissible that these bereversed, depending on the nature of the propellant, their densities andthe optimum mixture rate. For example, the oxidizer is preferablyprovided through the tangential inlets 62 to form the double vortex flowfield for the case where significantly more oxidizer is utilized thenfuel. Generally, depending on the particular fuel, about two times asmuch, to as much as six times as much, oxidizer is used as fuel.However, in some cases the ratios may be reversed and cause the designerto revise the propellant roles.

The embodiment of FIG. 6 is a further embodiment of a liquid rocketpropulsion system and includes a combustion chamber 71 defined by agenerally cylindrical chamber side wall 72 and a closed end defined bythe end wall 74. Positioned in the central portion of the end wall 74 isa fuel supply injector 75 comprising a plurality of fuel inlet ports oropenings 76. The fuel inlet ports or openings 76 are connected with theclosed end fuel manifold 78 which is in turn connected with the fuelsource inlet 79. In this embodiment, the fuel source is a fluid fuelsource which is as RP-1 (a kerosene type fuel). However, gaseous fuelsmay be used equally well.

The nozzle end of the chamber includes a converging nozzle housingportion 80 extending from the bottom end of the side wall 72 andconverging toward the nozzle throat 82. The nozzle end also includes adiverging nozzle portion 81 extending from the nozzle throat 82. Thelower or nozzle end of the side wall 72 is provided with a plurality ofoxidizer inlets 84 positioned around the periphery of the chamber. Theseare directed tangentially to the inner surface of the side wall 72. Anannular oxidizer manifold 85 encircles the side wall 72 in the area ofthe inlets 84 and is connected with a regenerative cooling skirt 86surrounding the converging portion 80 of the nozzle end. An oxidizerinlet 88 connects the cooling skirt 86 with an oxidizer source.

During operation, oxidizer is introduced into the inlet 88. From therethe oxidizer flows through the cooling skirt 86 through the oxidizerinlets 84 and tangentially into the chamber 71. As discussed previously,this forms a field flow comprising the outer vortex 89 and the innervortex 90. Unlike the other embodiments, the embodiment of FIG. 6introduces the fuel only at the closed end of the chamber from the fuelmanifold 78 through the fuel inlet ports 76. Likewise, ignition occursat the closed end of the chamber where the oxidizer is combined with theincoming fuel. Thus, in the embodiment of FIG. 6, all of the combustionoccurs in the inner vortex 90.

A significant advantage of this embodiment is the fact that the portionof the oxidizer flow in the outer vortex 89 spirals along the wall 72,without combustion, until it reaches the closed end, where it mixes withthe fuel and the mixture is ignited. Because no combustion can occur inthe outer vortex 89 along the wall 72 because of the absence of fuelthere, the wall 72 is kept relatively cool, despite the extremely hotcombustion temperatures in the inner vortex. This permits the combustionchamber to be constructed of cheaper materials which are less heatresistant than those of prior liquid rocket engines. It also greatlyincreases the service life of the chamber.

Although not specifically illustrated in the drawings, a controlmechanism is associated with each of the embodiments of FIGS. 3-6 tocontrol the flow of oxidizer into the system. Such control mechanismalso controls the supply of liquid or gaseous fuel in the embodiments ofFIGS. 5 and 6. Preferably, the flow of these fluids is controlled toprovide the desired combustion and thrust and to optimize the efficiencyto the extent possible. These control systems are well known to thoseproficient in the art and may be of any of several means and yet bewithin the spirit of this invention.

FIG. 7 is a schematic representation of a double stage rocket comprisinga rocket payload 91, a first liquid or hybrid stage 92 and a secondliquid or hybrid or liquid stage 94. The size of the stages 92 and 94may be varied to meet the particular propulsion requirements of therocket. Each of the first and second stages 92, 94 includes a combustionchamber 95 conforming substantially to one of the propulsion systemembodiments described above. Extending from the nozzle end of eachcombustion chamber 95 is an exhaust nozzle 96. Also associated with eachstage 92 and 94 is a fuel and oxidizer source 98. Positioned in asuitable compartment of the rocket is the rocket guidance and controlsystem 99.

FIG. 8 is a view of a further embodiment of a hybrid rocket engine. Thecombustion chamber 26 of FIG. 8 is defined in part by the solid fuelgrain 46. Further, the chamber 26 includes a wall portion with an innersurface and comprised of a side wall portion and an outlet end wallportion. The side wall portion includes the wall 44 and the fuel grain46, while the outlet end wall portion includes the base or end wall 30.This view is similar to that of FIG. 4 except that the inner wall 44 isprovided with inlet openings 47 at the lower end of the wall 44 adjacentto the base or outlet end wall 30. Preferably, there is no axialdistance between the upper surface of the base 30 and the inlets 47.Accordingly, when oxidizer or other fluid enters the combustion chamber26, all of the flow is directed in a tangential direction in a spiralpattern which flows upwardly toward the closed end wall 29. As shown,the embodiment of FIG. 8 is free of any inlet opening openings otherthan those adjacent to the wall 30.

A further embodiment of a hybrid rocket engine is shown in FIGS. 9 and10. In this embodiment, a plurality of oxidizer inlets 23 are providedin the base wall portion 20 as shown. Preferably these inlets 23 angleupwardly with an outward radial velocity component to ensure that theinjected fluid will travel in a spiral vortex up the wall 11 of thechamber 14. With specific reference to FIG. 9, the rocket engine isprovided with a nozzle portion 81 extending downwardly from the wallportion 20 and one or more inlet nozzles 27 forming a common manifoldfor supplying oxidizer or other fluid to the interior of the chamber 14through the inlets or injection ports 23. In the embodiment of FIGS. 9and 10, the side wall 11 is free of any inlet openings.

Although the descriptions of the preferred embodiments have been quitespecific, it is contemplated that various modifications could be madewithout deviating from the spirit of the present invention. Accordingly,it is intended that the present invention be dictated by the appendedclaims rather then by the descriptions of the preferred embodiments.

What is claimed is:
 1. A hybrid rocket engine comprising: a solid fuelgrain; a combustion chamber defined in part by said solid fuel grain,said combustion chamber having a substantially closed end, an outlet endhaving a fluid outlet opening, a wall portion between said closed endand said outlet opening, and an inner surface; and one or more fluidinlet openings in said wall portion for directing a fluid into saidchamber to create a spiral flow of said fluid along said inner surfacetoward said closed end.
 2. The rocket engine of claim 1 wherein saidwall portion includes a side wall portion and an outlet end wallportion.
 3. The rocket engine of claim 2 wherein at least one of saidone or more fluid inlet openings is positioned in said side wallportion.
 4. The rocket engine of claim 3 wherein at least one of saidone or more fluid inlet openings is positioned adjacent to said outletend wall portion.
 5. The rocket engine of claim 4 including a pluralityof said one or more fluid inlet openings positioned adjacent to saidoutlet end wall portion.
 6. The rocket engine of claim 5 wherein saidside wall portion is substantially free of said one or more fluid inletopenings other than said plurality of fluid inlet openings.
 7. Therocket engine of claim 6 wherein said side wall portion is generallyrevolute in geometry.
 8. The rocket engine of claim 1 wherein saidcombustion chamber is generally revolute in geometry.
 9. The rocketengine of claim 2 wherein at least one of said one or more fluid inletopenings is positioned in said outlet end wall portion.
 10. The rocketengine of claim 9 including a plurality of said one or more fluid inletopenings positioned in said outlet end wall portion.
 11. The rocketengine of claim 10 wherein said side wall portion is free of said one ormore fluid inlet openings.
 12. The rocket engine of claim 10 including acommon manifold for supplying said fluid to said plurality of fluidinlet openings.
 13. The rocket engine of claim 10 wherein said side wallportion is generally revolute in geometry.
 14. The rocket engine ofclaim 1 wherein said fuel grain forms a part of said walled portion. 15.A method of propelling a rocket comprising: providing a hybrid rocketengine having a combustion chamber defined in part by a fuel grain andhaving a substantially closed end, an outlet opening opposite saidclosed end, a wall portion between said closed end and said outletopening and one or more fluid inlets in said wall portion; introducing afluid comprising one of a fluid fuel and a fluid oxidizer into saidcombustion chamber through said fluid inlets to cause said fluid to movein a substantially spiral path along said wall portion and said fuelgrain toward said closed end to form a mixture of said fluid and saidfuel grain; and igniting said mixture in said combustion chamber. 16.The method of claim 15 including providing a combustion chamber in whichsaid wall portion includes a side wall portion and an outlet end wallportion.
 17. The method of claim 16 including providing a combustionchamber in which at least one of said fluid inlets is in said side wallportion adjacent to said outlet end wall portion.
 18. The method ofclaim 16 including providing a combustion chamber in which at least oneof said fluid inlets is in said outlet end wall portion.
 19. The methodof claim 16 including providing a combustion chamber in which said sidewall portion is substantially revolute in geometry.
 20. The method ofclaim 15 including providing a combustion chamber which is generallyrevolute in geometry.
 21. The method of claim 15 wherein said fluid is afluid oxidizer.
 22. A rocket engine comprising: a combustion chamberhaving a chamber axis and a wall portion encircling said chamber axis,said wall portion defining a closed end and a nozzle end and having aninner surface; one or more first fluid inlet openings in said wallportion for directing one of an oxidizer fluid and a fuel fluid intosaid chamber to create a spiral flow of said one fluid along said innersurface toward said closed end; and one or more second fluid inletopenings in said wall portion for introducing the other of said oxidizerfluid and said fuel fluid into said chamber.
 23. The rocket engine ofclaim 22 wherein said wall portion includes a side wall portion and aclosed end wall portion.
 24. The rocket engine of claim 23 wherein saidone or more second fluid inlet openings are in said closed end wallportion.
 25. The rocket engine of claim 23 wherein said one or moresecond fluid inlet openings are in said closed end wall portion and saidside wall portion.
 26. The rocket engine of claim 22 wherein said wallportion includes a nozzle end wall portion and said one or more firstfluid inlet openings are in said nozzle end wall portion.